Large gas turbine engines are widely used for aircraft propulsion and for ground based power generation. Such large gas turbine engines are of the axial type, and include a compressor section, a combustor section, and a turbine section, with the compressor section normally preceded by a fan section. An annular flow path for working medium gases extends axially through the sections of the engine. Each of the fan, compressor, and turbine sections comprises a plurality of disks mounted on a shaft, with a plurality of airfoil shaped blades projecting radially from the disks. A hollow case surrounds the various engine sections. A plurality of stationary vanes are located between the disks and project inwardly from the case assembly which surrounds the disks.
During operation of the fan, compressor, and turbine sections, as the working medium gases are flowed axially, they alternately contact moving blades and the stationary vanes. In the fan and compressor sections, air is compressed and the compressed air is combined with fuel and burned in the combustion section to provide high pressure, high temperature gases. The working medium gases then flow through the turbine section, where energy is extracted by causing the bladed turbine disks to rotate. A portion of this energy is used to operate the compressor section and the fan section.
Engine efficiency depends to a significant extent upon minimizing leakage of the gas flow to maximize interaction between the gas stream and the moving and stationary airfoils. A major source of inefficiency is leakage of gas around the tips of the compressor blades, between the blade tips and, the engine case. Accordingly, means to improve efficiency by reduction of leakage are increasingly important. Although a close tolerance fit may be obtained by fabricating the blade tips and the engine case to mate to a very close tolerance range, this fabrication process is extremely costly and time consuming. Further, when the assembly formed by mating the blade tips and the engine case is exposed to a high temperature environment and rotational forces, as when in use, the coefficients of expansion of the blade tips and the engine case parts may differ, thus causing the clearance space to either increase or decrease. A significant decrease in clearance results in contact between blades and housing, and friction between the parts generates heat causing a significant elevation of temperatures and possible damage to one or both members. On the other hand, increased clearance space would permit gas to escape between the compressor blade and housing, thus decreasing efficiency.
One approach to increase efficiency is to apply an abradable coating of suitable material to the interior surface of the compressor housing, which when abraded allows for the creation of a channel between the blade tips and the housing. Leakage between the blade tips and the housing is limited to airflow in the channel. Various coating techniques have been employed to coat the inside diameter of the compressor housing with an abradable coating that can be worn away by the frictional contact of the compressor blade, to provide a close fitting channel in which the blade tip may travel. Thus, when subjecting the coated assembly to a high temperature and stress environment, the blade and the case may expand or contract without resulting in significant gas leakage between the blade tip and the housing.
However, it is critical that the blade tips not degrade when contacted with the coatings applied to the interior surface of the compressor housing. To increase the durability of the blade tips which rub against the abradable seals, abrasive layers are sometimes applied to the blade tip surface.
The abrasive layers must have a particular combination of properties. They must be resistant to erosion from the high velocity, high temperature gas streams which at times may carry fine particulate matter with them. The abradable coating must also be structurally sound to resist the thermal and vibratory strains imposed upon it in use. In addition, the intentional contact between the abrasive tip and engine case creates a demanding, high wear environment for the abrasive blade tip coating.
Considerable effort has gone into the development of abradable coatings having the desired combination of properties. For example, Vine et al., U.S. Pat. No. 4,861,618 discloses a thermal barrier coating which may be used on the airfoil section of a turbine blade. In one embodiment, Vine et al. discloses a NiCoCrAlY bond coat with a ceramic overcoat of zirconia comprising six to eight weight percent (6 to 8 wt. %) yttria.
This above art notwithstanding, scientists and engineers working under the direction of Applicant's assignee are seeking to improve the composition of the abrasive coating applied to substrates in a gas turbine engine.